Turbine shroud sealing architecture

ABSTRACT

A shroud sealing arrangement for a gas turbine engine, which comprises a static shroud assembly mounted to an engine case and having a platform surrounding a rotatable airfoil array. The platform has an inner side and an outer side and extends from a leading edge to a trailing edge. A shroud support structure mounts the shroud platform to the case. A circumferential groove is defined on the outer side of the shroud platform proximal to one of the leading edge and the trailing edge. A sealing ring is set in the groove and adapted to seal cooling air from escaping directly to the gas path.

TECHNICAL FIELD

The application relates generally to gas turbine engines and, moreparticularly, to static shroud assemblies for rotor blade arrays.

BACKGROUND OF THE ART

Typically, an axial gap is provided between a turbine shroud and theouter wall of a gas path duct at ambient temperatures, to allow forthermal expansion of the duct and/or the turbine shroud at engineoperating temperatures. The magnitude of such thermal expansion can bepredicted, and the gap sized, so that thermal expansion generally sealsthe gap to prevent leakage through the gap.

However, the seal is not perfect and it must be ensured to adequatelypurge the adjacent cavity with sufficient cooling air to avoid hot gasingestion. Reducing such uses of secondary air can increase gas turbineengine efficiency.

Accordingly, there is a need for an improved turbine shroud sealingarrangement.

SUMMARY

In one aspect, there is provided a shroud sealing arrangement for a gasturbine engine, the arrangement comprising: a static shroud assemblymounted to an engine case and having a circumferential array of shroudsegments surrounding a rotatable blade array, the shroud segments eachhaving a platform, the platform having a radially inner side and aradially outer side and extending axially from a leading edge to atrailing edge, and a forward leg and an aft leg extending radiallyoutwardly from the radially outer side of the platform; a shroud supportstructure engaged with the forward and aft legs of the shroud segmentsfor mounting the shroud segments to the engine case; a circumferentiallyextending groove defined on the radially outer side of the shroudsegments proximal to one of the leading edge and the trailing edge; anda sealing ring mounted in the circumferentially extending groove, thesealing ring cooperating with the shroud support structure to define acooling air plenum with one of said forward and aft legs.

In another aspect, a gas turbine engine has a circumferential array ofshroud segments surrounding a rotatable blade array in a gas pathwhereby the shroud segments are secured to an engine case by a shroudsupport structure. An adjacent stator vane assembly forms a gap with thearray of shroud segments. An annular slot is defined in the shroudsegments near the gap and a radial sealing ring is set in the slot forsealing cooling air to the array of shroud segments.

In accordance with another aspect, there is provided a method forcooling the shroud segments of a circumferential array of shroudsegments surrounding a rotatable turbine blade array in a gas path, theshroud segments each having forward and aft legs extending radiallyoutwardly from a radially outer surface of a platform, the methodcomprising: capturing cooling air leaking from between the forward oraft legs in a cooling air plenum closing a leading edge or trailing edgecavity of the shroud segments, and reusing said cooling air to provideimpingement cooling on an adjacent component.

In accordance with a still further general aspect, there is provided amethod for cooling the shroud segments of a circumferential array ofshroud segments surrounding a rotatable turbine blade array in a gaspath, the method including: supplying cooling air to the array of shroudsegments, sealing the cooling air in the area of the shroud segments bydefining a radially outwardly facing annular slot near an edge of theshroud segments; providing a sealing ring in the slot and providingdischarge ports in the sealing ring

DESCRIPTION OF THE DRAWINGS

Reference is now made to the accompanying figures, in which:

FIG. 1 is a schematic, cross-sectional view of a turbofan engine havinga reverse flow annular combustor;

FIG. 2 is a schematic, fragmentary view in axial cross-section of theturbine shroud area of the engine shown in FIG. 1; and

FIG. 3 is schematic, fragmentary view in axial cross-section of theturbine shroud area similar to FIG. 2, but showing the cooling air flow.

DETAILED DESCRIPTION

FIG. 1 illustrates a gas turbine engine 10 of a type preferably providedfor use in subsonic flight, generally comprising in serial flowcommunication a fan 12 through which ambient air is propelled, amultistage compressor 14 for pressurizing the air, a combustor 16 inwhich the compressed air is mixed with fuel and ignited for generatingan annular stream of hot combustion gases, and a turbine section 18 forextracting energy from the combustion gases. A gas path 20 carries thehot combustion gases from the combustor through the turbine section 18for powering the turbines.

The combustor 16 is housed in a plenum 17 supplied with compressed airfrom compressor 14. The turbine section 18 is also surrounded by theplenum 17, defined within the engine case 22, for supplying cooling airto a turbine shroud surrounding the turbine blades 26 (see FIG. 2). Theturbine section 18 generally comprises one or more stages of turbineblades 26 extending radially outwardly from respective rotor disks, withthe blade tips 26 a being disposed closely adjacent to an annularturbine shroud 24 supported from the engine case 22. The shroud 24 istypically circumferentially segmented. FIGS. 2 and 3 illustrate anexample of one such turbine shroud segments 30. The various stages ofturbine blades 26 are arranged in the gas path 20 with alternatingstator vanes 28.

As seen in FIG. 2, each shroud segment 30 comprises axially spaced-apartforward and aft hooks or legs 32 and 40 extending radially outwardlyfrom a back side or cold radially outer surface 33 a of an arcuateplatform 33. The platform 33 has an opposite radially inner hot gas flowsurface 33 b adapted to be disposed adjacent to the tip 26 a of theturbine blades 26. The platform 33 is axially defined from a leadingedge 34 to a trailing edge 42 in a direction from an upstream positionto a downstream position of a hot gas flow passing through gas path 20,and being circumferentially and longitudinally defined between oppositelateral sides.

The forward leg 32 is disposed just downstream of the leading edge 34 ofthe platform 33. The leg 32 includes a fastener device 36, extending,axially downstream of the leg 32. The fastener device 36 engages ashroud support housing 38 mounted to the engine case 22.

The aft leg 40 is disposed upstream of the trailing edge 42 of theplatform 33. A projection 44 extends downstream and axially from the leg40. The projection 44 engages a corresponding axial recess 46 defined inthe shroud support housing 38. A cooling air chamber 48 is definedbetween the shroud support housing 38 and the forward and aft legs 32,40 of the shrouds segments 30. Bores 50 traverse the shroud supporthousing 38 and communicate the plenum 17 with the cooling air chamber48.

Axial gaps 52 are typically provided between the stator shroud 54 andthe leading edge 34 of the shroud segments 30 to provide for thermalexpansion. Cooling air can escape through the gaps 52 to exhaust intothe gas path 20.

A circumferentially extending slot or groove 58 is defined in theradially outer surface 33 a of the platform 33 of the shroud segments 30axially between the leading edge 34 and the forward leg 32. The grooves58 of the shroud segments 30 collectively form a full or 360 degreesgroove. A 360 degrees sealing ring 56 is mounted in the fullcircumferential groove 58 formed by the shroud segments 30. The sealingring 56 may be provided in the form of a lightweight, annular metalplate.

As shown in FIG. 2, the outer portion 56 a, of sealing ring 56, mayaxially contact the sealing surface 38 a of the shroud support housing38. A circumferential W seal 68 is also resilient and adds pressure tothe annular ring 56 to engage the seal surface 38 a. An axial, contactsealing surface 60 is defined on a short axial stub 62 which projectsupstream from the annular ring 56 radially inwardly from the outer orperipheral portion 56 a. Part of the stator shroud aft support leg 55includes a contact surface 64 defined on a short axial stub 66 opposedto the contact surface 60. Surfaces 60 and 64 form contact sealing facesin running conditions.

Referring now to FIG. 3, which is identical to FIG. 2, there is shown byway of arrows the movement of the cooling air emanating from the plenum17. The cooling air enters the shroud array 24 through the bores 50 inthe shroud support housing 38 to the cooling air chamber 48. As there isno feather seal on the forward legs 32 of the shroud segments 30, theair, under pressure, within the cooling air chamber 48 will leak throughthe interface between adjacent forward legs 32 of the shroud segments30. This leakage air is received in a cooling air plenum 72 definedbetween the annular ring 56 and the forward leg 32 of the shroudsegments 30. The air in plenum 72 provides cooling along all the lengthof the forward leg 32. It also provides for a better cooling of theleading edge region of the platform. This contributes to improve shrouddurability. It also eliminates the need for multiple feather sealsbetween the forward legs of the shroud segments. Air also passes by theaft legs 40 in order to enter the plenum 49 where the cooling air canimpinge on the downstream portion of the platform of the shroud segments30. Along the axial length of the platform 33 of the shroud segments 30are feather seals 76 and cooling air impinges on the shroud segment 30,between the feather seals 76.

Cooling air passes from the plenum 72 through impingement holes 70defined in the sealing ring 56. The holes 70 may be evenly distributedon a circumferential row and oriented so as to aim at the back face ofthe adjacent stator shroud 54. The size and number of discharge ports orholes will be determined by design criteria for a given engine. Asdepicted by the arrows in FIG. 3, the air passing through the holes 70impinges on the back face of the stator shroud 54. The air may then beused to purge the gap 52 formed between the stator shroud 54 and theannular ring 56 as well as the leading edge 34 of the shroud segments30.

Reusing the cooling air to cool the adjacent component (the statorshroud) and to purge the gap between the shroud segments and theadjacent component allows to reduce the amount of cooling air and, thuscontributes to the engine efficiency. The 360 degrees sealing platearchitecture also provides better control of cooling air leakage ascompared to individual feather seals.

During operation, the hot environment of the gas path 20 causes theshroud segments 30 and the stator vane shroud 54 as well as shroudsupport 55 to expand axially towards each other so that the contactsurfaces 60 and 64 of the stubs 62 and 66 respectively sealingly engageeach other, thus providing a seal against the loss of the cooling airinto the gas path 20. At the same time, the W seal 68 is compressed sothat the outer portion 56 a of the sealing ring 56 abuts the contactsurface 38 a in a sealing arrangement. However a nominal amount ofcooling air loss is acceptable. The spent cooling air once into the gaspath 20 may form a cooling film along the outer surface of the shroudsegments 30.

The above description is meant to be exemplary only, and one skilled inthe art will recognize that changes may be made to the embodimentdescribed without departing from the scope of the invention disclosed.For example, the sealing ring 56 can be provided with differentconfigurations, and is not limited to application in turbofan engines.Furthermore the spring shown in the drawings can have differentconfigurations and need only be resilient. Also, the sealing ring couldbe mounted in an associated groove defined in the radially outer surfaceof the platform axially between the aft leg and the trailing edge of theplatform to provide sealing along the aft leg and ensure proper coolingthereof. Still other modifications which fall within the scope of thepresent invention will be apparent to those skilled in the art, in lightof a review of this disclosure, and such modifications are intended tofall within the appended claims.

What is claim is:
 1. A shroud sealing arrangement for a gas turbineengine, the arrangement comprising: a static shroud assembly mounted toan engine case and having a circumferential array of shroud segmentssurrounding a rotatable blade array, the shroud segments each having aplatform, the platform having a radially inner side and a radially outerside and extending axially from a leading edge to a trailing edge, and aforward leg and an aft leg extending radially outwardly from theradially outer side of the platform; a shroud support structure engagedwith the forward and aft legs of the shroud segments for mounting theshroud segments to the engine case; a circumferentially extending groovedefined on the radially outer side of the shroud segments proximal toone of the leading edge and the trailing edge; and a sealing ringmounted in the circumferentially extending groove, the sealing ringcooperating with the shroud support structure to define a cooling airplenum with one of said forward and aft legs.
 2. The shroud sealingarrangement defined in claim 1, wherein air passages are defined in theshroud support structure to direct cooling air in a cooling chamberdefined between the forward and aft legs of the shroud segments, andwherein the cooling air plenum defined by the sealing ring is in fluidflow communication with the cooling chamber.
 3. The shroud sealingarrangement defined in claim 1, wherein impingement holes are defined inthe sealing ring, the impingement holes aiming at an adjacent structureto direct impingement jets thereagainst.
 4. The shroud sealingarrangement as defined in claim 1, wherein the sealing ring has an axialsealing face at a peripheral outer end thereof, the axial sealing facebeing in sealing engagement with a corresponding sealing face of theshroud support structure.
 5. The shroud sealing arrangement as definedin claim 2, wherein the static shroud assembly is adjacent a stator vaneassembly, upstream of the shroud assembly and an axial gap is formedtherebetween, the sealing ring being mounted between the leading edgeand the forward leg, and wherein the sealing ring is provided with holesallowing cooling air from the cooling chamber to be reused to purge theaxial gap.
 6. The shroud sealing arrangement as defined in claim 1,wherein the shroud support structure has a circumferentially extendingfront sealing contact surface and the sealing ring has a cylindricalstub extending downstream thereof and having a contact surface adaptedto engage the circumferentially extending front sealing contact surfaceof the shroud support structure to seal the cooling air within theshroud assembly.
 7. The shroud sealing arrangement as defined in claim5, wherein the stator vane assembly has a rearwardly axially facingsealing contact surface and the sealing ring has a cylindrical stubextending upstream thereof and having a corresponding sealing contactsurface adapted to engage the rearwardly axially facing sealing contactsurface of the stator vane assembly to seal the outer area of the axialgap when the engine is in operation and the thermal, axial expansion ofthe shroud assembly and the stator vane assembly has caused the axialgap to be reduced.
 8. The shroud sealing arrangement as defined in claim5, wherein a resilient, circumferential seal is provided between thestator vane assembly and the sealing ring at an outer radial portionthereof in order to bias the sealing ring in sealing contact with theshroud support structure.
 9. In a gas turbine engine having acircumferential array of shroud segments surrounding a rotatable bladearray in a gas path; the shroud segments secured to an engine case by ashroud support structure, an adjacent stator vane assembly forming a gapwith the array of shroud segments; air passages in the shroud supportstructure to allow cooling air to contact the shroud segments; anannular slot is defined in the shroud segments adjacent the gap, and asealing ring is set in the slot for sealing cooling air to the array ofshroud segments.
 10. In the gas turbine engine as defined in claim 9,wherein the shroud support structure has a circumferential sealingcontact surface and the sealing ring includes a cylindrical stubextending downstream thereof and having a contact surface adapted toengage the circumferential sealing contact surface of the shroud supportstructure to seal the cooling air within the shroud assembly; andwherein the stator vane assembly has an annular sealing contact surfaceand the sealing ring includes a cylindrical stub having a correspondingcontact surface adapted to engage the annular sealing contact surface ofthe stator vane assembly to seal the outer area of the gap when theengine is in operation and the thermal, axial expansion of the shroudsegments and the stator vane assembly has caused the gap to be reduced.11. In the gas turbine engine as defined in claim 10, the sealing ringis provided with impingement holes allowing cooling air to be dischargedinto the gap to impinge on the stator vane assembly and exhausting tothe gas path.
 12. A method for cooling the shroud segments of acircumferential array of shroud segments surrounding a rotatable turbineblade array in a gas path, the shroud segments each having forward andaft legs extending radially outwardly from a radially outer surface of aplatform, the method comprising: capturing cooling air leaking frombetween the forward or aft legs in a cooling air plenum closing aleading edge or trailing edge cavity of the shroud segments, and reusingsaid cooling air to provide impingement cooling on an adjacentcomponent.
 13. The method defined in claim 12, further comprisingreusing the cooling air after impingement cooling to purge a gap betweenthe adjacent component and the shroud segments.
 14. The method definedin claim 12, wherein reusing said cooling air to provide impingementcooling comprises causing the cooling air to flow through impingementholes aiming at said adjacent component.
 15. The method defined in claim14, wherein the capturing step comprises mounting a sealing ring in acircumferential groove defined in a radially outer surface of the shroudsegments, the sealing ring being in sealing contact with a shroudsupport structure to which the shroud segments are mounted, and whereinthe impingement holes are defined in the sealing ring.